1. Field of the Art
The present invention relates generally to a vortex flow field and the apparatus and method to produce and sustain it. The flow field in accordance with the present invention is capable of providing separate regions or zones within and among one or more flowing fluids contained within a common chamber, without the need for diaphragms or other physical separators or barriers. It is evident and believed that the flow field of the present invention has utility to a wide range of applications. One general field of application is that of combustion chambers, and more particularly, that of combustion chambers and methods for rocket engines or the like. A combustion chamber and method in accordance with one embodiment of the present invention utilizes the unique flow field of the present invention to improve hybrid rocket fuel regression and increase mixing length in a rocket or other engine. Another embodiment is in the form of liquid rocket engine to prevent hot combustion products from contacting the chamber wall.
2. Description of the Prior Art
Virtually countless applications exist for a flow field which is compact and is capable of providing one or more separate regions or zones of flowing fluids within a container, without substantial mixing and without the need for any physical barrier or other separators between such regions or zones. With such a flow field, a chemical reaction, such as combustion, can be induced to incur in one region or zone while a separate fluid or process occupies another region or zone. Accordingly, such a flow field is useful in connection with the design of various types of combustion chambers and methods, as well as various types of separators and other devices described as follows.
Many devices depend upon vortex flows for their successful operation, such as combustion chambers, cyclone separators, classifiers and the like that are in common use. All of these devices introduce swirling flow at one end of a passageway in which the flow follows a generally helical path to exit at the opposite end.
Although the flow field in accordance with the present invention has significant applications in a variety of fields, it has particular application to the field of rocket engines and al in one embodiment specifically to hybrid rocket engines. In recent years, hybrid rockets have received increasing attention from National Aeronautics and Space Administration (NASA) sectors, Department of Defense, industrial aerospace participants and research institutions because their unique operational characteristics are capable of providing safer, lower-cost avenues to space then conventional solid propellant and liquid bi-propellant rocket propulsion systems. For example, hybrid rocket engines can be easily throttled for thrust tailoring, to perform in-flight motor shutdown and restart and to incorporate non-destructive mission abort modes. Also, since fuel in a hybrid rocket engine is stored in the form of a solid grain, such engines require only half the feed system hardware of liquid bi-propellant engines. Still further, the commonly used butadiene-based solid grain fuels are benign and neither toxic nor hazardous for storage and transportation. The hybrid solid fuel grain is also not susceptible to cracks and imperfections as are solid rocket motor propellant grains.
However, despite these benefits, classical hybrid rocket engines, in which the oxidizer gas is injected into the combustion chamber at the end opposite the exit nozzle and in a direction parallel to the solid fuel grain, have not yet found widespread use for either commercial or military applications. Reasons for this include the fact that they suffer from relatively slow solid fuel regression rates, low volumetric loading and relatively poor combustion efficiency. For example, polymeric hybrid fuels such as hydroxyl-terminated polybutadiene (HTPD) regress generally about an order of magnitude slower than solid rocket motor propellants. In an effort to overcome these lower regression rates, complex cross-sectional geometries of the hybrid solid grain fuel with large wetted surface areas are often employed to achieve a large mass of flow rate of pyrolyzed vapor from the fuel grain. It has been shown that a three to fourfold increase in fuel regression rate can result in significant cost reductions, simplified grain manufacturing and large reductions in rocket inert weight.
In addition to problems associated with the low regression (fuel burning) rates of hybrid engines, the short straight line travel of the pyrolyzed fuel grain vapor and oxidizer result in incomplete mixing and often necessitate the use of secondary combustion chambers at the end of the fuel grain to complete the combustion. These chambers add length and weight to the overall design and have the additional disadvantage of serving as a potential source and location of combustion instability.
Accordingly, there is a need in the art for a flow field, and a structure and method for producing and sustaining it, which provides separate regions or zones of flowing fluids within a chamber. There is also a need in the art for a combustion chamber and method utilizing such a flow field, and particularly a combustion chamber and method for a hybrid rocket engine, which significantly increases the regression rate of the solid fuel grain and the effective chamber length and mixing within the combustion chamber.